The subject matter disclosed herein relates generally to gas turbine engines, and, more particularly, to a film-cooling augmentation device and turbine airfoil that incorporates the same.
Gas turbine engines generally include a compressor that pressurizes air that is then mixed with fuel in a combustor for generating hot combustion gases. Energy from these combustion gases is extracted in multiple turbine stages for powering the compressor and producing useful work, typically by powering a fan in an aircraft turbofan application, or by powering an output shaft for marine, industrial and other applications.
The combustion gases flow along and transfer heat to various components of the turbine engine. These heated components are cooled in part by a cooling system that uses a portion of the pressurized air generated by the compressor in a cooling circuit that extends to the surface of these components to provide film-cooling of their surfaces that are exposed to the hot combustion gases, or that are heated thereby. The film cooled turbine components include the combustor, nozzles, rotor disks, blades, vanes, shrouds, liners and other components.
The hot combustion gases are initially discharged from the combustor into a stationary high pressure turbine inlet comprising the first stage turbine nozzle which includes an array of circumferentially-spaced, radially-extending stator nozzle vanes that include vane airfoils which extend between inner and outer bands. The vane airfoils are partially hollow and form a part of the cooling circuit and include various rows and other patterns of film-cooling holes extending through their sidewalls for discharging film-cooling air to form a boundary layer over their outer surfaces, particularly at the leading and trailing edges of the airfoils. The other stages of the turbine also include similar nozzles and arrays of vanes.
The first stage of the turbine also includes a turbine rotor including an array of circumferentially-spaced, radially-extending turbine rotor blades that are attached to an outer portion of a supporting rotor disk. Each blade includes a partially hollow dovetail portion that is attached to a partially hollow turbine airfoil where their hollow portions form a part of the cooling circuit and include various rows and other patterns of film-cooling holes extending through the sidewalls thereof for film-cooling the outer surfaces thereof, particularly at the leading and trailing edges of the airfoil. The other stages of the turbine also include similar rotors disks and arrays of vanes.
Film-cooling holes are also found in other components of the typical gas turbine engine and are arranged in various patterns for promoting a film-cooling blanket of air over the surfaces that are exposed either directly or indirectly to the hot combustion gases. The film-cooling holes are frequently arranged in linear rows, with the rows being spaced laterally apart for distributing the film-cooling air as required for accommodating the local heat loads from the combustion gases and forming the blanket of cooling air.
A typical film-cooling hole is tubular or cylindrical and manufactured by laser drilling for example. Another form of film-cooling hole is the diffusion hole which has various configurations in the art. In the diffusion hole the outlet portion thereof diverges or increases in flow area in the downstream aft direction from the upstream inlet for reducing the discharge velocity therefrom. An exemplary diffusion hole has a trapezoidal outlet with side edges which diverge at a suitably small diffusion angle, and an inner land which blends with the component outboard surface at a shallower inclination angle than the nominal inclination angle through the inlet portion of the hole. In this way, the typical diffusion hole is effective for laterally spreading the discharged cooling air jet and locally enhancing film-cooling performance.
The configuration, quantity, and pattern of the film-cooling holes, including their cross-sectional shape, length and angle with the surface to be cooled are specifically designed for the expected heat load, which varies from component to component and over the outboard surface of an individual component. Generally, film-cooling hole design, including size, shape, pattern, location and other aspects, has as an objective reduction of the amount of film-cooling air bled from the compressor, since use of such air for film-cooling prevents its use in the combustion process and thus reduces efficiency of the engine.
The performance of film-cooling holes is affected by many factors, including their geometry including their surface area, and the local conditions in the specific components including the differential pressure or pressure ratio between the outboard and inboard sides of the film-cooling holes, and the velocity and pressure distribution of the combustion gases over the outboard surfaces.
While film-cooling holes, including diffusion holes, are useful to provide film-cooling in the manner noted above, the ability to control the pressure differential or pressure ratio between the outboard and inboard sides of the film-cooling holes and the airflow through them, as well as the amount of convective cooling of the airfoil sidewall adjacent to the hole have generally been limited to control of the characteristics of the holes themselves, such as their location, spacing, size, shape, orientation (e.g., orientation of the hole inlet to the cooling circuit and angle with the airfoil surface), number and the like, which are in turn have been limited by the manufacturing methods used to create them and other considerations, such as the overall performance requirements of the airfoil.
Accordingly, it is desired to provide improved methods of cooling gas turbine engine components from the film-cooling holes utilized therein.